Efficient, low pressure ratio propulsor for gas turbine engines

ABSTRACT

A gas turbine engine includes a gear assembly, a bypass flow passage, a fan located upstream of the bypass flow passage, a first shaft and a second shaft, a first turbine coupled through the gear assembly to the fan, a first compressor coupled with the first shaft, and a second turbine coupled with the second shaft. The fan has a bypass ratio of greater than 8.5. The fan includes a hub and a row of fan blades that extend from the hub. The row includes a number (N) of the fan blades that is from 16 to 20, a solidity value (R) at tips of the fan blades, and a ratio of N/R that is from 12.3 to 20.

CROSS REFERENCE TO RELATED APPLICATIONS

The present disclosure is a continuation of U.S. application Ser. No. 16/515,506, filed Jul. 18, 2019, which is a continuation of U.S. application Ser. No. 15/709,567, filed Sep. 20, 2017, which is a continuation of U.S. application Ser. No. 14/695,373, filed Apr. 24, 2015, which is a continuation-in-part of U.S. application Ser. No. 13/484,858, filed May 31, 2012, which is a continuation of U.S. application Ser. No. 13/176,365, filed Jul. 5, 2011.

STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT

This invention was made with government support under contract number NAS3-01138 awarded by NASA. The government has certain rights in the invention.

BACKGROUND

This disclosure relates to gas turbine engines and, more particularly, to an engine having a geared turbo fan architecture that is designed to efficiently operate with a high bypass ratio and a low pressure ratio.

The overall propulsive efficiency and fuel burn of a gas turbine engine depends on many different factors, such as the design of the engine and the resulting performance debits on the fan that propels the engine. As an example, the fan rotates at a high rate of speed such that air passes over the blades at transonic or supersonic speeds. The fast-moving air creates flow discontinuities or shocks that result in irreversible propulsive losses. Additionally, physical interaction between the fan and the air causes downstream turbulence and further losses. Although some basic principles behind such losses are understood, identifying and changing appropriate design factors to reduce such losses for a given engine architecture has proven to be a complex and elusive task.

SUMMARY

A gas turbine engine according to an example of the present disclosure includes a gear assembly, a bypass flow passage, a fan located upstream of the bypass flow passage and that has a bypass ratio of greater than 8.5, a first shaft and a second shaft, a first turbine coupled through the gear assembly to the fan, and a first compressor coupled with the first shaft. The first compressor is a 3-stage compressor. A second turbine is coupled with the second shaft. The fan includes a hub and a row of fan blades that extend from the hub, and the row includes a number (N) of the fan blades that is from 16 to 20, a solidity value (R) at tips of the fan blades, and a ratio of N/R that is from 12.3 to 20.

In a further embodiment of any of the foregoing embodiments, the second turbine is a 2-stage turbine.

In a further embodiment of any of the foregoing embodiments, the bypass flow passage includes an inlet and an outlet defining a design pressure ratio with regard to an inlet pressure at the inlet and an outlet pressure at the outlet at a design rotational speed of the engine, and the design pressure ratio is less than 1.55.

In a further embodiment of any of the foregoing embodiments, the bypass flow passage includes an inlet and an outlet defining a design pressure ratio with regard to an inlet pressure at the inlet and an outlet pressure at the outlet at a design rotational speed of the engine, and the design pressure ratio is from 1.3 to 1.4.

A further embodiment of any of the foregoing embodiments includes a case surrounding the fan, and the fan is a single fan stage.

In a further embodiment of any of the foregoing embodiments, each of the fan blades is fixed in position between the hub and the tip.

In a further embodiment of any of the foregoing embodiments, the fan blades are formed of a fiber-reinforced polymer matrix material.

In a further embodiment of any of the foregoing embodiments, the first turbine is a 5-stage turbine.

A gas turbine engine according to an example of the present disclosure includes a gear assembly, a bypass flow passage, a fan located upstream of the bypass flow passage and that has a bypass ratio of greater than 8.5, a first shaft and a second shaft, a first turbine coupled through the gear assembly to the fan, a first compressor coupled with the first shaft, and a second turbine coupled with the second shaft. The second turbine is a 2-stage turbine. The fan includes a hub and a row of fan blades that extend from the hub. The row includes a number (N) of the fan blades that is from 16 to 20, a solidity value (R) at tips of the fan blades, and a ratio of N/R that is from 12.3 to 20.

In a further embodiment of any of the foregoing embodiments, the bypass flow passage includes an inlet and an outlet defining a design pressure ratio with regard to an inlet pressure at the inlet and an outlet pressure at the outlet at a design rotational speed of the engine, and the design pressure ratio is less than 1.55.

In a further embodiment of any of the foregoing embodiments, the bypass flow passage includes an inlet and an outlet defining a design pressure ratio with regard to an inlet pressure at the inlet and an outlet pressure at the outlet at a design rotational speed of the engine, and the design pressure ratio is from 1.3 to 1.4.

A further embodiment of any of the foregoing embodiments includes a case surrounding the fan, and the fan is a single fan stage.

In a further embodiment of any of the foregoing embodiments, each of the fan blades is fixed in position between the hub and the tip.

In a further embodiment of any of the foregoing embodiments, the fan blades are formed of a fiber-reinforced polymer matrix material.

In a further embodiment of any of the foregoing embodiments, the first turbine is a 5-stage turbine.

In a further embodiment of any of the foregoing embodiments, the first compressor is a 3-stage compressor.

A gas turbine engine according to an example of the present disclosure includes a gear assembly, a bypass flow passage, a fan located upstream of the bypass flow passage and that has a bypass ratio of greater than 8.5, a first shaft and a second shaft, and a first turbine coupled through the gear assembly to the fan. The first turbine is a 5-stage turbine. A first compressor is coupled with the first shaft. A second turbine is coupled with the second shaft. The fan includes a hub and a row of fan blades that extend from the hub, and the row includes a number (N) of the fan blades that is from 16 to 20, a solidity value (R) at tips of the fan blades, and a ratio of N/R that is from 12.3 to 20.

In a further embodiment of any of the foregoing embodiments, the first compressor is a 3-stage compressor.

In a further embodiment of any of the foregoing embodiments, the second turbine is a 2-stage turbine.

In a further embodiment of any of the foregoing embodiments, the bypass flow passage includes an inlet and an outlet defining a design pressure ratio with regard to an inlet pressure at the inlet and an outlet pressure at the outlet at a design rotational speed of the engine, and the design pressure ratio is less than 1.55.

In a further embodiment of any of the foregoing embodiments, the bypass flow passage includes an inlet and an outlet defining a design pressure ratio with regard to an inlet pressure at the inlet and an outlet pressure at the outlet at a design rotational speed of the engine, and the design pressure ratio is from 1.3 to 1.4.

A further embodiment of any of the foregoing embodiments includes a case surrounding the fan, and the fan is a single fan stage.

In a further embodiment of any of the foregoing embodiments, each of the fan blades is fixed in position between the hub and the tip.

In a further embodiment of any of the foregoing embodiments, the fan blades are formed of a fiber-reinforced polymer matrix material.

BRIEF DESCRIPTION OF THE DRAWINGS

The various features and advantages of the disclosed examples will become apparent to those skilled in the art from the following detailed description. The drawings that accompany the detailed description can be briefly described as follows.

FIG. 1 is a schematic cross-section of an embodiment of a gas turbine engine.

FIG. 2 is a perspective view of a fan section of the engine of FIG. 1.

FIG. 3 illustrates an embodiment of a carbon-fiber reinforced polymer matrix material.

FIG. 4 illustrates an embodiment of a two-dimensional woven fiber structure.

FIG. 5 illustrates an embodiment of a three-dimensional fiber structure.

FIG. 6 is a cross-section of an embodiment of a propulsor blade that has a distinct core and a skin of carbon-fiber reinforced polymer matrix material.

FIG. 7 illustrates an embodiment of a propulsor blade that has a sheath.

FIG. 8 illustrates a portion of an embodiment of a case and propulsor blade.

FIG. 9 illustrates another embodiment of a case.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT

FIG. 1 schematically illustrates a gas turbine engine 20. The gas turbine engine 20 may be a two-spool turbofan that generally incorporates a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28. Alternative engine architectures may include a single-spool design, a three-spool design, or an open rotor design, among other systems or features.

The fan section 22 drives air along a bypass flow passage B while the compressor section 24 drives air along a core flow passage C for compression and communication into the combustor section 26. Although depicted as a turbofan gas turbine engine, it is to be understood that the concepts described herein are not limited to use with turbofans and the teachings may be applied to other types of gas turbine engines.

The engine 20 includes a low speed spool 30 and high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. The low speed spool 30 generally includes an inner shaft 40 that is coupled with a propulsor 42, a low pressure compressor 44 and a low pressure turbine 46. The propulsor 42 is in the fan section 22 and a case 43 surrounds the propulsor 42. The low pressure turbine 46 drives the propulsor 42 through the inner shaft 40 and a gear assembly 48, which allows the low speed spool 30 to drive the propulsor 42 at a different (e.g. lower) angular speed.

The high speed spool 32 includes an outer shaft 50 that is coupled with a high pressure compressor 52 and a high pressure turbine 54. A combustor 56 is arranged between the high pressure compressor 52 and the high pressure turbine 54. The inner shaft 40 and the outer shaft 50 are concentric and rotate about the engine central longitudinal axis A, which is collinear with their longitudinal axes.

A core airflow in core flow passage C is compressed by the low pressure compressor 44 then the high pressure compressor 52, mixed with the fuel in the combustor 56, and then expanded over the high pressure turbine 54 and low pressure turbine 46. The turbines 54, 46 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion.

As shown, the propulsor 42 is arranged at an inlet 60 of the bypass flow passage B and core flow passage C. Air flow through the bypass flow passage B exits the engine 20 through an outlet 62 or nozzle. For a given design of the propulsor 42, the inlet 60 and the outlet 62 of the engine 20 define a design (fan) pressure ratio with regard to an inlet pressure at the inlet 60 and an outlet pressure at the outlet 62 of the bypass flow passage B. As an example, the design pressure ratio may be determined based upon the stagnation inlet pressure and the stagnation outlet pressure at a design rotational speed of the engine 20. In that regard, the engine 20 may optionally include a variable area nozzle 64 within the bypass flow passage B. The variable area nozzle 64 is operative to change a cross-sectional area 66 of the outlet 62 to thereby control the pressure ratio via changing pressure within the bypass flow passage B. The design pressure ratio may be defined with the variable area nozzle 64 fully open or fully closed.

Referring to FIG. 2, the propulsor 42, which in this example is a fan, includes a rotor 70 having a row 72 of propulsor blades 74 that extend a circumferentially around a hub 76. Each of the propulsor blades 74 extends radially outwardly from the hub 76 between a root 78 and a tip 80 and in a chord direction (axially and circumferentially) between a leading edge 82 and a trailing edge 84. A chord dimension (CD) is a length between the leading edge 82 and the trailing edge 84 at the tip of each propulsor blade 74. The row 72 of propulsor blades 74 also defines a circumferential pitch (CP) that is equivalent to the arc distance between the tips 80 of neighboring propulsor blades 74.

As will be described, the example propulsor 42 includes a number (N) of the propulsor blades 74 and a geometry that, in combination with the architecture of the engine 20, provides enhanced overall propulsive efficiency by reducing performance debits of the propulsor 42.

In the illustrated example, the number N of propulsor blades in the row 72 is no more than 20. In one example, the propulsor 42 includes 18 of the propulsor blades 74 uniformly circumferentially arranged about the hub 76. In other embodiments, the number N may be any number of blades from 12-20.

The propulsor blades 74 define a solidity value with regard to the chord dimension CD and the circumferential pitch CP. The solidity value is defined as a ratio (R) of CD/CP (i.e., CD divided by CP). In embodiments, the solidity value of the propulsor 42 is between 0.9 or 1.0 and 1.3. In further embodiments, the solidity value is from 1.1 to 1.2. In additional embodiments, the solidity value is less than 1.1, and in a further example is also greater than 0.85.

Additionally, in combination with the given example solidity values, the fan 22 of the engine 20 may be designed with a particular design pressure ratio. In embodiments, the design pressure ratio may be between 1.2 or 1.3 and 1.55. In a further embodiment, the design pressure ratio may be between 1.3 and 1.4. In further examples, the design pressure ratio is between 1.3 and 1.7.

The engine 20 may also be designed with a particular bypass ratio with regard to the amount of air that passes through the bypass flow passage B and the amount of air that passes through the core flow passage C. As an example, the design bypass ratio of the engine 20 may nominally be 12, or alternatively in a range of approximately 8.5 to 13.5 or 18.

The propulsor 42 also defines a ratio of N/R. In embodiments, the ratio N/R is from 9 to 20. In further embodiments, the ratio N/R is from 14 to 16. The table below shows additional examples of solidity and the ratio N/R for different numbers of propulsor blades 74.

TABLE Number of Blades, Solidity and Ratio N/R Number of Blades Ratio (N) Solidity N/R 20 1.3 15.4 18 1.3 13.8 16 1.3 12.3 14 1.3 10.8 12 1.3 9.2 20 1.2 16.7 18 1.2 15.0 16 1.2 13.3 14 1.2 11.7 12 1.2 10.0 20 1.1 18.2 18 1.1 16.4 16 1.1 14.5 14 1.1 12.7 12 1.1 10.9 20 1.0 20.0 18 1.0 18.0 16 1.0 16.0 14 1.0 14.0 12 1.0 12.0

The disclosed ratios of N/R enhance the overall propulsive efficiency and fuel burn of the disclosed engine 20. For instance, the disclosed ratios of N/R are designed for the geared turbo fan architecture of the engine 20 that utilizes the gear assembly 48. That is, the gear assembly 48 allows the propulsor 42 to rotate at a different, lower speed than the low speed spool 30. In combination with the variable area nozzle 64, the propulsor 42 can be designed with a large diameter and rotate at a relatively slow speed with regard to the low speed spool 30. A relatively low speed, relatively large diameter, and the geometry that permits the disclosed ratios of N/R contribute to the reduction of performance debits, such as by lowering the speed of the air or fluid that passes over the propulsor blades 74.

The propulsor blades 74 can include a carbon-fiber reinforced polymer matrix material, an example portion of which is depicted in FIG. 3 at 86. In this example, the material 86 includes carbon fibers 86 a that are disposed in a polymer matrix 86 b. The propulsor blades 74 can be formed exclusively of the material 86 or partially of the material 86 in combinations with alloys or other fiber-reinforced materials.

The material 86 can include a plurality of carbon fiber layers 88 that are stacked and consolidated to form the material 86. For example, the fiber layers 88 can each have uni-directionally oriented fibers and the layers 88 can be cross-plied. In further examples, one or more of the layers 88 has a different fiber structure, such as but not limited to, random fiber orientation, woven, or three-dimensional. An example two-dimensional woven fiber structure is depicted in FIG. 4. An example three-dimensional fiber structure is depicted in FIG. 5. In this example, the fibers 86 a are woven into sheets 90, and transverse fibers 86 c bundle the sheets 90 to one another. As can be appreciated, other two- or three-dimensional fiber structures could alternatively or additionally be used.

The polymer matrix 86 b can include thermoplastic polymer, thermoset polymer, or combinations thereof. Thermoset polymers can include, but are not limited to, epoxy and phenolic. Thermoplastic polymers can include, but are not limited to, polyethers and polyimides.

The carbon fibers 86 a provide the material 86 with strength and stiffness. For example, the properties of the carbon fibers 86 a can be selected in accordance with desired properties of the material 86, and thus desired properties of the propulsor blades 74. In one example, the carbon fibers 86 a are polyacrylonitrile or polyacrylonitrile-based. The fibers are initially with polyacrylonitrile fibers and are then graphitized. Alternatively, the fibers are initially thermoplastic fibers that are then graphitized. Thermoplastics can include, but are not limited to, polyethylene, polyarylether, and poly ether ketones. In further examples, the carbon fibers 86 a have an average diameter of 1-100 micrometers. Alternatively, the carbon fibers 86 a are nano-sized and have a diameter of less than 1 micrometer. In other examples, the carbon fibers 86 a are carbon-containing such that the fibers include carbon as a primary constituent or element. In one example, the carbon fibers 86 a are carbide.

FIG. 6 illustrates a cross-sectional view of another example propulsor blade 174, which may include any of the aforementioned features. In this disclosure, like reference numerals designate like elements where appropriate and reference numerals with the addition of one-hundred or multiples thereof designate modified elements that are understood to incorporate the same features and benefits of the corresponding elements. In this example, the propulsor blade 174 includes a distinct core 174 a that supports a skin 174 b of the carbon-fiber reinforced polymer matrix material 86. In this example, the core 174 a is a solid piece, but it alternatively can be hollow to reduce weight.

The core 174 a can be formed of a metallic material, a fiber reinforced polymer matrix material, or combinations thereof. An example metallic material includes a titanium-based alloy. The fiber reinforced polymer matrix material can include carbon fiber, as in any of the examples of the material 86. Alternatively, the fibers in the core 174 a are non-carbon fibers. Example non-carbon fibers can include, but are not limited to, glass fibers, metallic fibers, ceramic fibers, polymeric fibers, and combinations thereof.

In further examples, the core 174 is formed of a fiber-reinforced material that is different in composition from the material 86 of the skin 174 b. The difference in composition can be in the kinds of polymers of the matrices, the kinds of fibers, the amounts of the polymer matrices, the amounts of the fibers, or any combination of such differences.

In further examples, the skin 174 b is the multi-layered structure of the material 86. For example, layers 88 are laid-up on or around the core 174 a and then consolidated. Alternatively, the skin 174 b is a continuous sleeve. The core 174 a is inserted into the sleeve and then the skin 174 b is consolidated. In one further example, the material 86 of the sleeve has a three-dimensional fiber structure.

FIG. 7 illustrates another example propulsor blade 274 that is formed of the material 86. In this example, the propulsor blade 274 also includes a sheath 275 on a leading edge of the blade. For example, the sheath 275 protects the propulsor blade 274 from foreign object impact. In one example, the sheath 275 is formed of a metallic material. The metallic material can include, but is not limited to, a titanium-based alloy, a cobalt-based alloy, or combinations thereof. In further examples, the sheath 275 is multi-layered and includes at least one layer of a metallic material. One or more additional layers can include a layer of a metallic material of a different composition, a layer of a polymer-based material, or combinations thereof.

The sheath 275 is secured to the leading edge of the propulsor blade 274. In this regard, the sheath 275 can be bonded using an adhesive, mechanically attached to the blade, or secured by a combination of adhesive bonding and mechanical attachment.

In a further example, the propulsor blade 274 includes a first distinct region 289 a (outside of dashed line region) of carbon-fiber reinforced polymer matrix material 86 and a second distinct region 289 b (inside dashed line region) of a non-carbon-fiber reinforced polymer matrix material. The non-carbon fibers can include, but are not limited to, glass fibers, aramid fibers, boron fibers, carbide fibers, or combinations thereof. The second distinct region 289 b of non-carbon-fiber reinforced polymer matrix material provides the ability to locally tailor the performance of the propulsor blade 274 with regard to properties. For example, the vibrational properties are locally tailored through selection of the properties of the second distinct region 289 b to control vibration or control response to an impact event.

FIG. 8 illustrates selected portions of the fan section 22 of the engine 20, including the case 43 and a portion of one of the propulsor blades 74. The case 43 serves as a containment structure in the case of a blade release event. For example, the case 43 includes a fiber reinforced polymer matrix material 45. The material 45 includes fibers 45 a that are disposed in a polymer matrix 45 b. The fibers 45 a can be carbon fibers or non-carbon fibers. Non-carbon fibers can include, but are not limited to, glass fibers, aramid fibers, or combinations thereof. In one example, the material 45 includes a plurality of fiber layers 45 c that are stacked and consolidated to form the material 45. For example, all of the layers 45 c have the same kind of fibers. In other examples, alternating layers 45 c, or an alternating pattern of layers 45 c, have different kinds of fibers, one of which is carbon fibers.

In further examples, the carbon-fiber reinforced polymer matrix material 86 of the propulsor blades 74 is different from the carbon-fiber reinforced polymer matrix material 45 of the case 43 with respect to composition. The difference in composition can be in the kinds of polymers of the matrices, the kinds of fibers, the amounts of the polymer matrices, the amounts of the fibers, or any combination of such differences. Further, the differences can be tailored for thermal conformance between the propulsor blades 74 and the case 43.

FIG. 9 illustrates another example case 143 that includes a layer of the material 45 adjacent a layer 147. The layer 147 can be a layer of carbon-fiber reinforced polymer matrix material, non-carbon-fiber reinforced polymer matrix material, or metallic material, such as in a honeycomb or acoustic structure.

Although a combination of features is shown in the illustrated examples, not all of them need to be combined to realize the benefits of various embodiments of this disclosure. In other words, a system designed according to an embodiment of this disclosure will not necessarily include all of the features shown in any one of the Figures or all of the portions schematically shown in the Figures. Moreover, selected features of one example embodiment may be combined with selected features of other example embodiments.

The preceding description is exemplary rather than limiting in nature. Variations and modifications to the disclosed examples may become apparent to those skilled in the art that do not necessarily depart from the essence of this disclosure. The scope of legal protection given to this disclosure can only be determined by studying the following claims. 

What is claimed is:
 1. A gas turbine engine comprising: a gear assembly; a bypass flow passage; a fan located upstream of the bypass flow passage and having a bypass ratio of greater than 8.5; a first shaft and a second shaft; a first turbine coupled through the gear assembly to the fan; a first compressor coupled with the first shaft, wherein the first compressor is a 3-stage compressor; a second turbine coupled with the second shaft; and wherein the fan includes a hub and a row of fan blades that extend from the hub, and the row includes a number (N) of the fan blades that is from 16 to 20, a solidity value (R) at tips of the fan blades, and a ratio of N/R that is from 12.3 to
 20. 2. The gas turbine engine as recited in claim 1, wherein the second turbine is a 2-stage turbine.
 3. The gas turbine engine as recited in claim 2, wherein the bypass flow passage includes an inlet and an outlet defining a design pressure ratio with regard to an inlet pressure at the inlet and an outlet pressure at the outlet at a design rotational speed of the engine, and the design pressure ratio is less than 1.55.
 4. The gas turbine engine as recited in claim 2, wherein the bypass flow passage includes an inlet and an outlet defining a design pressure ratio with regard to an inlet pressure at the inlet and an outlet pressure at the outlet at a design rotational speed of the engine, and the design pressure ratio is from 1.3 to 1.4.
 5. The gas turbine engine as recited in claim 3, further comprising a case surrounding the fan, and the fan is a single fan stage.
 6. The gas turbine engine as recited in claim 5, wherein each of the fan blades is fixed in position between the hub and the tip.
 7. The gas turbine engine as recited in claim 6, wherein the fan blades are formed of a fiber-reinforced polymer matrix material.
 8. The gas turbine engine as recited in claim 6, wherein the first turbine is a 5-stage turbine.
 9. A gas turbine engine comprising: a gear assembly; a bypass flow passage; a fan located upstream of the bypass flow passage and having a bypass ratio of greater than 8.5; a first shaft and a second shaft; a first turbine coupled through the gear assembly to the fan; a first compressor coupled with the first shaft; a second turbine coupled with the second shaft, wherein the second turbine is a 2-stage turbine; and wherein the fan includes a hub and a row of fan blades that extend from the hub, and the row includes a number (N) of the fan blades that is from 16 to 20, a solidity value (R) at tips of the fan blades, and a ratio of N/R that is from 12.3 to
 20. 10. The gas turbine engine as recited in claim 9, wherein the bypass flow passage includes an inlet and an outlet defining a design pressure ratio with regard to an inlet pressure at the inlet and an outlet pressure at the outlet at a design rotational speed of the engine, and the design pressure ratio is less than 1.55.
 11. The gas turbine engine as recited in claim 9, wherein the bypass flow passage includes an inlet and an outlet defining a design pressure ratio with regard to an inlet pressure at the inlet and an outlet pressure at the outlet at a design rotational speed of the engine, and the design pressure ratio is from 1.3 to 1.4.
 12. The gas turbine engine as recited in claim 10, further comprising a case surrounding the fan, and the fan is a single fan stage.
 13. The gas turbine engine as recited in claim 12, wherein each of the fan blades is fixed in position between the hub and the tip.
 14. The gas turbine engine as recited in claim 13, wherein the fan blades are formed of a fiber-reinforced polymer matrix material.
 15. The gas turbine engine as recited in claim 14, wherein the first turbine is a 5-stage turbine.
 16. The gas turbine engine as recited in claim 13, wherein the first compressor is a 3-stage compressor.
 17. A gas turbine engine comprising: a gear assembly; a bypass flow passage; a fan located upstream of the bypass flow passage and having a bypass ratio of greater than 8.5; a first shaft and a second shaft; a first turbine coupled through the gear assembly to the fan, wherein the first turbine is a 5-stage turbine; a first compressor coupled with the first shaft; a second turbine coupled with the second shaft; and wherein the fan includes a hub and a row of fan blades that extend from the hub, and the row includes a number (N) of the fan blades that is from 16 to 20, a solidity value (R) at tips of the fan blades, and a ratio of N/R that is from 12.3 to
 20. 18. The gas turbine engine as recited in claim 17, wherein the first compressor is a 3-stage compressor.
 19. The gas turbine engine as recited in claim 18, wherein the second turbine is a 2-stage turbine.
 20. The gas turbine engine as recited in claim 19, wherein the bypass flow passage includes an inlet and an outlet defining a design pressure ratio with regard to an inlet pressure at the inlet and an outlet pressure at the outlet at a design rotational speed of the engine, and the design pressure ratio is less than 1.55.
 21. The gas turbine engine as recited in claim 19, wherein the bypass flow passage includes an inlet and an outlet defining a design pressure ratio with regard to an inlet pressure at the inlet and an outlet pressure at the outlet at a design rotational speed of the engine, and the design pressure ratio is from 1.3 to 1.4.
 22. The gas turbine engine as recited in claim 20, further comprising a case surrounding the fan, and the fan is a single fan stage.
 23. The gas turbine engine as recited in claim 22, wherein each of the fan blades is fixed in position between the hub and the tip.
 24. The gas turbine engine as recited in 23, wherein the fan blades are formed of a fiber-reinforced polymer matrix material. 